Mechanically Joining Airframe Members at Solid Insert

ABSTRACT

An airframe assembly for an aircraft includes a first airframe member having first and second skins with a large cell core and a solid insert joined therebetween. The solid insert has a side surface at least a portion of which is adjacent to the large cell core. The first airframe member has a first set of openings extending through the first skin, the solid insert and the second skin. The airframe assembly also includes a second airframe member having a second set of openings operable to be aligned with the first set of openings of the first airframe member. Each of a plurality of fasteners extends through one of the openings of the first set of openings and one of the openings of the second set of openings securably coupling the first airframe member to the second airframe member.

CROSS-REFERENCE TO RELATED APPLICATIONS

This patent application is a continuation-in-part of copending U.S.application Ser. No. 15/424,095, filed Feb. 3, 2017, which claims thebenefit of the filing date of U.S. Provisional Patent Application No.62/292,718, filed Feb. 8, 2016, both of which are hereby incorporated byreference.

TECHNICAL FIELD OF THE DISCLOSURE

The present disclosure relates, in general, to airframe members formedfrom lightweight, high stiffness structural panels and, in particular,to airframe members formed from large cell core stiffened panels havingsolid inserts at load points that provide joining locations for couplingthe airframe members with other airframe structures.

BACKGROUND

The airframes of modern aircraft are constructed from a wide variety ofmaterials, including steel, aluminum and composites. While most airframecomponents are made from strong, rigid materials, in order to conserveweight, certain airframe components are made from relatively thinmaterial layers attached to stiffening structures such as stringers. Forexample, the wing of a conventional tiltrotor aircraft includes a torquebox structure formed from an upper skin, a lower skin, a forward sparand an aft spar. The upper and lower skins have stringers attachedthereto that extend generally parallel with the longitudinal axis of thewing to provide stiffness and support to the skins. The stringers mayhave an I-beam cross section and are typically connected to the interiorsurface of the skins at reinforcement strips that provide support forthe skins against catastrophic buckling, help to maintain the shape andcontour of the skins, provide stiffness at the stringer load points anddistribute pressure into the skins. In addition, the torque boxstructure typically includes multiple internal support members thatprovide horizontal structural strength to the forward and aft spars andthe upper and lower skins.

It has been found, however, that the assembly of the torque boxstructure for conventional tiltrotor aircraft wings is complex andrequires very tight tolerances. For example, the installation ofnumerous fasteners to the skins and other structural components isdifficult and time consuming due to limited access to small interiorspaces and complicated sealing requirements. Also, once the structuralmembers are assembled, numerous foam details must be positioned betweenthe structural members in the fuel bays to provide a smooth, rampedsurface for the fuel components housed therein. In addition, it has beenfound, that the thickness of stringers as well as the multiple internalsupport members reduce the space available for fuel and other internalsystems within the torque box structure.

SUMMARY

In a first aspect, the present disclosure is directed to a corestiffened panel that includes a first skin and a second skin. A largecell core is joined between the first and second skins. A solid inserthaving a side surface is also joined between the first and second skinssuch that at least a portion of the side surface is adjacent to thelarge cell core.

In some embodiments, the large cell core is joined to the adjacentportion of the side surface of the solid insert. In certain embodiments,the large cell core may be formed from an array of cells having a widthof between about 0.5 inches and about 1.5 inches and a thickness ofbetween about 0.25 inches and about 0.75 inches. In some embodiments,the large cell core may be a large cell composite core such as a largecell carbon core. In certain embodiments the first skin, the second skinand the solid insert may be formed from a composite materials such as acarbon composite material such that the first skin, the second skin andthe solid insert have generally matching coefficients of thermalexpansion.

In some embodiments, the large cell core and the solid insert may bestructurally bonded between the first and second skins and/or the largecell core may be structurally bonded to the adjacent portion of the sidesurface of the solid insert. In certain embodiments, the solid insertextends between first and second ends of at least one of the first andsecond skins. In other embodiments, the solid insert may be surroundedby the large cell core. In some embodiments, a plurality of solidinserts may be joined between and/or structurally bonded to the firstand second skins. In certain embodiments, the solid insert may be asymmetrical solid insert, a non symmetrical solid insert and/or a solidinsert with a void therein. In certain embodiments, the first skin, thesolid insert and the second skin may include at least one openingextending therethrough.

In a second aspect, the present disclosure is directed to a method offorming a core stiffened panel. The method includes providing a firstskin; disposing a large cell core on the first skin; locating a solidinsert having a side surface on the first skin such that at least aportion of the side surface is adjacent to the large cell core;positioning a second skin on the large cell core and the solid insertopposite the first skin to form a panel assembly; and curing the panelassembly to join the large cell core to the first and second skins andto join the solid insert to the first and second skins, thereby formingthe core stiffened panel.

The method may also include joining the large cell core to the solidinsert; structurally bonding the large cell core to the first and secondskins; structurally bonding the solid insert to the first and secondskins; surrounding the solid insert with the large cell core; locating aplurality of solid inserts on the first skin and joining each of thesolid inserts to the first and second skins and/or forming at least oneopening extending through the first skin, the solid insert and thesecond skin.

In a third aspect, the present disclosure is directed to an airframeassembly for an aircraft. The airframe assembly includes a firstairframe member having a first skin, a second skin, a large cell corejoined between the first and second skins and a solid insert having aside surface. The solid insert is joined between the first and secondskins such that at least a portion of the side surface is adjacent tothe large cell core. The first skin has a first surface disposedopposite the solid insert. The airframe assembly also includes a secondairframe member having a second surface. An adhesive joint is disposedbetween the first and second surfaces structurally bonding the firstairframe member to the second airframe member such that the secondairframe member is positioned opposite the solid insert.

In a fourth aspect, the present disclosure is directed to a wingassembly for an aircraft. The wing assembly includes a wing skin havingan inner skin member, an outer skin member, a large cell core joinedbetween the inner and outer skin members and a solid insert having aside surface. The solid insert is joined between the inner and outerskin members such that at least a portion of the side surface isadjacent to the large cell core. The inner skin member has a firstsurface disposed opposite the solid insert. The wing assembly alsoincludes a rib having a second surface. An adhesive joint is disposedbetween the first and second surfaces structurally bonding the rib tothe inner skin member of the wing skin such that the rib is positionedopposite the solid insert.

In a fifth aspect, the present disclosure is directed to an aircraft.The aircraft includes a first airframe member having a first skin, asecond skin, a large cell core joined between the first and second skinsand a solid insert having a side surface. The solid insert is joinedbetween the first and second skins such that at least a portion of theside surface is adjacent to the large cell core. The first skin has afirst surface disposed opposite the solid insert. The airframe assemblyalso includes a second airframe member having a second surface. Anadhesive joint is disposed between the first and second surfacesstructurally bonding the first airframe member to the second airframemember such that the second airframe member is positioned opposite thesolid insert.

In a sixth aspect, the present disclosure is directed to an airframeassembly for an aircraft. The airframe assembly includes a firstairframe member having a first skin, a second skin, a large cell corejoined between the first and second skins and a solid insert having aside surface. The solid insert is joined between the first and secondskins such that at least a portion of the side surface is adjacent tothe large cell core. The first airframe member has a first set ofopenings extending through the first skin, the solid insert and thesecond skin. The airframe assembly also includes a second airframemember having a second set of openings operable to be aligned with thefirst set of openings of the first airframe member. Each of a pluralityof fasteners extends through one of the openings of the first set ofopenings and one of the openings of the second set of openings securablycoupling the first airframe member to the second airframe member.

In a seventh aspect, the present disclosure is directed to a wingassembly for an aircraft. The wing assembly includes a wing skin havingan inner skin member, an outer skin member, a large cell core joinedbetween the inner and outer skin members and a solid insert having aside surface. The solid insert is joined between the inner and outerskin members such that at least a portion of the side surface isadjacent to the large cell core. The wing skin has a first set ofopenings extending through the inner skin member, the solid inset andthe outer skin member. The wing assembly also includes a rib having asecond set of openings operable to be aligned with the first set ofopenings of the wing skin. Each of a plurality of fasteners extendsthrough one of the openings of the first set of openings and one of theopenings of the second set of openings securably coupling the rib to thewing skin.

In an eighth aspect, the present disclosure is directed to an aircraft.The aircraft includes a first airframe member having a first skin, asecond skin, a large cell core joined between the first and second skinsand a solid insert having a side surface. The solid insert is joinedbetween the first and second skins such that at least a portion of theside surface is adjacent to the large cell core. The first airframemember has a first set of openings extending through the first skin, thesolid insert and the second skin. The airframe assembly also includes asecond airframe member having a second set of openings operable to bealigned with the first set of openings of the first airframe member.Each of a plurality of fasteners extends through one of the openings ofthe first set of openings and one of the openings of the second set ofopenings securably coupling the first airframe member to the secondairframe member.

BRIEF DESCRIPTION OF THE DRAWINGS

For a more complete understanding of the features and advantages of thepresent disclosure, reference is now made to the detailed descriptionalong with the accompanying figures in which corresponding numerals inthe different figures refer to corresponding parts and in which:

FIGS. 1A-1B are schematic illustrations of a tiltrotor aircraft in aforward flight mode and in a vertical takeoff and landing flight mode,respectively, in accordance with embodiments of the present disclosure;

FIG. 2 is an exploded view of a torque box structure of a wing of atiltrotor aircraft in accordance with embodiments of the presentdisclosure;

FIG. 3 is a flow diagram of a method of manufacturing a large cell corestiffened panel having solid inserts in accordance with embodiments ofthe present disclosure;

FIGS. 4A-4F are schematic illustrations of a method of manufacturing alarge cell core stiffened panel having solid inserts in accordance withembodiments of the present disclosure;

FIGS. 5A-5B are top and cross sectional views of a large cell core foruse in a large cell core stiffened panel having solid inserts inaccordance with embodiments of the present disclosure;

FIGS. 6A-6B are schematic illustrations of a wing rib joined to a wingskin formed from a large cell core stiffened panel having solid insertsin accordance with embodiments of the present disclosure;

FIGS. 7A-7B are schematic illustrations of a large cell core stiffenedpanel having solid inserts in accordance with embodiments of the presentdisclosure;

FIGS. 8A-8B are schematic illustrations of a large cell core stiffenedpanel having solid inserts in accordance with embodiments of the presentdisclosure;

FIGS. 9A-9B are schematic illustrations of a large cell core stiffenedpanel having solid inserts in accordance with embodiments of the presentdisclosure;

FIGS. 10A-10B are schematic illustrations of a large cell core stiffenedpanel having solid inserts in accordance with embodiments of the presentdisclosure; and

FIGS. 11A-11B are schematic illustrations of a large cell core stiffenedpanel having solid inserts in accordance with embodiments of the presentdisclosure.

DETAILED DESCRIPTION

While the making and using of various embodiments of the presentdisclosure are discussed in detail below, it should be appreciated thatthe present disclosure provides many applicable inventive concepts,which can be embodied in a wide variety of specific contexts. Thespecific embodiments discussed herein are merely illustrative and do notdelimit the scope of the present disclosure. In the interest of clarity,not all features of an actual implementation may be described in thepresent disclosure. It will of course be appreciated that in thedevelopment of any such actual embodiment, numerousimplementation-specific decisions must be made to achieve thedeveloper's specific goals, such as compliance with system-related andbusiness-related constraints, which will vary from one implementation toanother. Moreover, it will be appreciated that such a development effortmight be complex and time-consuming but would be a routine undertakingfor those of ordinary skill in the art having the benefit of thisdisclosure.

In the specification, reference may be made to the spatial relationshipsbetween various components and to the spatial orientation of variousaspects of components as the devices are depicted in the attacheddrawings. However, as will be recognized by those skilled in the artafter a complete reading of the present disclosure, the devices,members, apparatuses, and the like described herein may be positioned inany desired orientation. Thus, the use of terms such as “above,”“below,” “upper,” “lower” or other like terms to describe a spatialrelationship between various components or to describe the spatialorientation of aspects of such components should be understood todescribe a relative relationship between the components or a spatialorientation of aspects of such components, respectively, as the devicedescribed herein may be oriented in any desired direction.

Referring to FIGS. 1A-1B in the drawings, a tiltrotor aircraft isschematically illustrated and generally designated 10. Aircraft 10includes a fuselage 12, a wing mount assembly 14 that is rotatablerelative to fuselage 12 and a tail assembly 16 having control surfacesoperable for horizontal and/or vertical stabilization during forwardflight. A wing 18 is supported by wing mount assembly 14 and rotateswith wing mount assembly 14 relative to fuselage 12 to enable tiltrotoraircraft 10 convert to a storage configuration. Together, fuselage 12,tail assembly 16 and wing 18 as well as their various frames, longerons,stringers, bulkheads, spars, ribs and skins such as aft wing spar 18 a,wing ribs 18 b and upper wing skin 18 c, may be considered to be theairframe of tiltrotor aircraft 10. In the illustrated embodiment, thewing skins include large cell core stiffened panels having solid insertsat various load points that provide joining locations for wing ribs 18b.

Located proximate the outboard ends of wing 18 are fixed nacelles 20 a,20 b, each of which preferably houses an engine and a fixed portion of adrive system. A pylon assembly 22 a is rotatable relative to fixednacelle 20 a and wing 18 between a generally horizontal orientation, asbest seen in FIG. 1A, a generally vertical orientation, as best seen inFIG. 1B. Pylon assembly 22 a includes a rotatable portion of the drivesystem and a proprotor system 24 a that is rotatable responsive totorque and rotational energy provided via the engine and drive system.Likewise, a pylon assembly 22 b is rotatable relative to fixed nacelle20 b and wing 18 between a generally vertical orientation, as best seenin FIG. 1A, a generally horizontal orientation, as best seen in FIG. 1B.Pylon assembly 22 b includes a rotatable portion of the drive system anda proprotor system 24 b that is rotatable responsive to torque androtational energy provided via the engine and drive system. In theillustrated embodiment, proprotor systems 24 a, 24 b each include fourproprotor blades 26. It should be understood by those having ordinaryskill in the art, however, that proprotor assemblies 24 a, 24 b couldalternatively have a different number of proprotor blades, either lessthan or greater than four. In addition, it should be understood that theposition of pylon assemblies 22 a, 22 b, the angular velocity orrevolutions per minute (RPM) of the proprotor systems 24 a, 24 b, thepitch of proprotor blades 26 and the like are controlled by the pilot oftiltrotor aircraft 10 and/or the flight control system to selectivelycontrol the direction, thrust and lift of tiltrotor aircraft 10 duringflight.

FIG. 1A illustrates tiltrotor aircraft 10 in a forward flight mode orairplane flight mode, in which proprotor systems 24 a, 24 b arepositioned to rotate in a substantially vertical plane to provide aforward thrust while a lifting force is supplied by wing 18 such thattiltrotor aircraft 10 flies much like a conventional propeller drivenaircraft. FIG. 1B illustrates tiltrotor aircraft 10 in a verticaltakeoff and landing (VTOL) flight mode or helicopter flight mode, inwhich proprotor systems 24 a, 24 b are positioned to rotate in asubstantially horizontal plane to provide a vertical thrust such thattiltrotor aircraft 10 flies much like a conventional helicopter. Duringoperation, tiltrotor aircraft 10 may convert from helicopter flight modeto airplane flight mode following vertical takeoff and/or hover.Likewise, tiltrotor aircraft 10 may convert back to helicopter flightmode from airplane flight mode for hover and/or vertical landing. Inaddition, tiltrotor aircraft 10 can perform certain flight maneuverswith proprotor systems 24 a, 24 b positioned between airplane flightmode and helicopter flight mode, which can be referred to as conversionflight mode.

Preferably, each fixed nacelle 20 a, 20 b houses a drive system, such asan engine and transmission, for supplying torque and rotational energyto a respective proprotor system 24 a, 24 b. In such embodiments, thedrive systems of each fixed nacelle 20 a, 20 b may be coupled togethervia one or more drive shafts located in wing 18 such that either drivesystem can serve as a backup to the other drive system in the event of afailure. Alternatively or additionally, fuselage 12 may include a drivesystem, such as an engine and transmission, for providing torque androtational energy to each proprotor system 24 a, 24 b via one or moredrive shafts located in wing 18. In tiltrotor aircraft having bothnacelle and fuselage mounted drive systems, the fuselage mounted drivesystem may serve as a backup drive system in the event of failure ofeither or both of the nacelle mounted drive systems.

Referring next to FIG. 2 of the drawings, an exploded view of a torquebox structure of wing 18 of tiltrotor aircraft 10 is depicted. Torquebox structure 50 includes an aft spar 52, a forward spar 54, a lowerskin assembly 56 including an outer skin member and an inner skin memberwith a large cell core and solid inserts joined therebetween, an upperskin assembly 58 including an outer skin member and an inner skin memberwith a large cell core and solid inserts joined therebetween, and aplurality of ribs 60. The various torque box components are may joinedtogether by adhesive bonding or using aerospace fasteners such as pins,screws, rivets or other suitable fastening means to form torque boxstructure 50.

Various structural components of the an airframe such the outer andinner skin members, spars, ribs and the like may be formed fromcomposite materials that may include numerous material plies composed ofcontinuous filaments or fibers including one or more of glass, carbon,graphite, basalt, aromatic polyamide materials or the like and anycombination thereof. The material plies may be in the form of fabricssuch as woven fabrics, tape such as unidirectional tape and the like.The plies may be joined together with a resin such as a polymeric matrixincluding thermoplastic or thermosetting resin or any suitable resinsystem such as epoxies, polyimides, polyamides, bismaleimides,polyesters, vinyl esters, phenolics, polyetheretherketones (PEEK),polyetherketones (PEK), polyphenylene sulfides (PPS) and the like. Eventhough the methods of manufacture, the composites, the components andthe subassemblies thereof are described herein in the context of atiltrotor aircraft, it should be understood by those having ordinaryskill in the art that the methods of manufacture, the composites, thecomponents and the subassemblies thereof can be implemented on otheraircraft including manned and unmanned aircraft.

In one example, the various airframe components may be formed using alayup process wherein a plurality of plies, which may be preimpregnatedwith an uncured resin, are placed in a mold or other support structure.The plies may have the same or different shapes, may be the same ordifferent materials, may have the same or different fabric weaves, mayhave the same or different thicknesses, may be continuous ordiscontinuous, may extend beyond the periphery of the mold or be fullycontained within the mold, may be oriented in the same or differentdirections and/or may have other similarities or differences. In oneembodiment, the plies may be an intermediate modulus epoxy resinimpregnated carbon fiber fabric that is stiffer than conventionalcomposite fabrics, thereby allowing for fewer plies and reducing theweight and manufacturing cost. Each fabric layer is considered a ply ofthe laminate with the total number of plies ranging from between about 8plies to about 100 plies for an outer skin member of a wing skin. Inother airframe implementations, the total number of plies may be betweenabout 10 plies to about 20 plies, between about 20 plies to about 30plies, between about 20 plies to about 60 plies, between about 40 pliesto about 60 plies and/or other suitable range or number of plies.

Following layup of the plies, the plies may be compacted to remove anypockets of air and to provide adequate attachment therebetween. Thecompacting step can be achieved by applying a vacuum to the plies in themold or by pressing a second mold member on the plies. After the pliesare compacted, the plies may be cured to form a laminate by heating theplies for two hours at 350 degrees Fahrenheit and at 90 psi, forexample. Following the curing process, the laminate may be trimmed orotherwise machined as desired. This and/or similar processes may be usedto form various components or subassemblies thereof for torque boxstructure 50 includes aft spar assembly 52, forward spar assembly 54,ribs 60 and/or the outer and inner skin members of lower skin assembly56 and upper skin assembly 58.

Referring next to FIGS. 3 and 4A-4F, manufacturing steps associated withforming a lower wing skin assembly for a torque box structure oftiltrotor aircraft 10 will now be described. In step 102 of method 100,a first skin 202, as best seen in FIG. 4A, is placed in a mold or othersecure structure (not pictured). First skin 202 is preferably a laminateformed from a plurality of material plies as discussed herein. In step104, a plurality of large cell core sections 204 are provided. Largecell core sections 204 may be in the form of large cell honeycombstructures, wherein the term “honeycomb” means a material comprising aplurality of interconnected cell walls that define a plurality of cells.The cells may take the form of hexagonal cells, rectangular cells,square cells, flex-core cells, reinforced cells or the like. The term“large cell,” for the purposes of this disclosure, means that each cell206 has a width W of at least 0.5 inches (see FIG. 5A). For example, thewidth W of the large cells may be between about 0.5 inches and about 1.5inches and preferably about 1 inch. The thickness T of the large cellsmay be between about 0.25 inches and about 0.75 inches and preferablyabout 0.5 inches (see FIG. 5B). Even though particular width andthickness dimensions have been described for the cells of large cellcore sections 204, those having ordinary skill in the art will recognizethat cells having other dimensions both larger and smaller than thosedescribed are possible and are considered to be within the scope of thepresent disclosure.

Large cell core sections 204 may be formed from a variety of materials,including but not limited to, composite materials and metals. Forexample, the walls of large cell core sections 204 may be made from oneor more material plies oriented in one or more directions and can bewoven, unwoven or braided, for example. Large cell core sections 204 maybe made of resin impregnated filaments or fibers composed of one or moreof glass, carbon, graphite, basalt, aromatic polyamide materials or thelike and any combination thereof. The resin may be a polymeric matrixincluding thermoplastic or thermosetting resin or any suitable resinsystem such as epoxies, polyimides, polyamides, bismaleimides,polyesters, vinyl esters, phenolics, polyetheretherketones (PEEK),polyetherketones (PEK), polyphenylene sulfides (PPS) and the like. Eachlarge cell core section 204 may be cut or machined to any suitable sizeor shape including the illustrated rectangular shape. In the illustratedembodiment, each large cell core section 204 is constructed of similarmaterial, shape and size but, in other embodiments, the plurality oflarge cell core sections 204 may include a variety of large cell coresections 204 having at least one different material, shape or size ascompared to the other large cell core sections 204.

In step 106, the plurality of large cell core sections 204 are disposedon first skin 202, as best seen in FIG. 4B. In the illustratedembodiment, large cell core sections 204 are positioned generally alongthe longitudinal axis of first skin 202 within the periphery of firstskin 202. In other embodiments, a single large cell core section 204 mayextend spanwise across the entire length of first skin 202.

In step 108, a plurality of solid inserts 208 are located on first skin202 such that at least a portion of a side surface of each solid insert208 is adjacent to at least one of the large cell core sections 204, asbest seen in FIG. 4C. Solid inserts 208 may be formed from a variety ofmaterials, including but not limited to, composite materials and metals.For example, solid inserts 208 may include numerous material pliescomposed of continuous filaments or fibers including one or more ofglass, carbon, graphite, basalt, aromatic polyamide materials or thelike and any combination thereof. Solid inserts 208 may include materialplies in the form of fabrics such as woven fabrics, tape such asunidirectional tape and the like. The plies may be joined together witha resin such as a polymeric matrix including thermoplastic orthermosetting resin or any suitable resin system such as epoxies,polyimides, polyamides, bismaleimides, polyesters, vinyl esters,phenolics, polyetheretherketones (PEEK), polyetherketones (PEK),polyphenylene sulfides (PPS) and the like. Each solid insert 208 may becut or machined to any suitable size or shape including the illustratedrectangular shape. Each solid insert 208 may alternatively be acompression molded material that is molded in the desired shape. Forexample, the compression molded material can be a curable moldablematerial, such as thermosetting resins and advanced compositethermoplastics with unidirectional tapes, woven fabrics, randomlyorientated fiber mat or chopped strand. Solid inserts 208 may be buttjointed to the adjacent large cell core sections 204 using a foamingadhesive or other suitable jointing method, if desired. Regardless ofthe selected materials or processes, first skin 202, solid inserts 208and second skin 210 preferably have generally matching coefficients ofthermal expansion.

In step 110, second skin 210 is positioned onto solid inserts 208 andlarge cell core sections 204, as best seen in FIG. 4D. In theillustrated embodiment, second skin 210 is a laminate formed from thesame materials and processes as first skin 202. In other embodiments,however, second skin 210 could be formed from other materials and byother processes. In the illustrated embodiment, second skin 210 isaligned with the outer edges of solid inserts 208 and large cell coresections 204. In other embodiments, second skin 210 could have adifferent length and/or width then solid inserts 208 and/or large cellcore sections 204.

In step 112, one or more adhesives or adhesive layers are applied to theupper surface of large cell core sections 204, the lower surface oflarge cell core sections 204, the upper surface of solid inserts 208,the lower surface of solid inserts 208, the upper surface of first skin202 and/or the lower surface of second skin 210. This step may takeplace at any suitable interval or intervals between step 102 and step110 such that the adhesive is disposed between the lower surface oflarge cell core sections 204 and first skin 202, the lower surface ofsolid inserts 208 and first skin 202, the upper surface of large cellcore sections 204 and second skin 210 as well as the upper surface ofsolid inserts 208 and second skin 210. In the illustrated embodiment,the adhesive may be reticulated to provide fillets of adhesive betweenfirst skin 202, large cell core sections 204 and second skin 210. Atthis stage, first skin 202, large cell core sections 204, solid inserts208 and second skin 210 may be referred to as a panel assembly 212and/or an uncured panel assembly.

In step 114, panel assembly 212, which may be represent by FIG. 4D, iscured under heat and pressure. For example, panel assembly 212 may becured for about 2 hours at 250 degrees Fahrenheit at a pressure aboveambient pressure, depending upon the adhesive or adhesives being cured.Following the curing step, large cell core sections 204 are joinedbetween first skin 202 and second skin 210 and are preferablystructurally bonded between first skin 202 and second skin 210. Inaddition, solid inserts 208 are joined between first skin 202 and secondskin 210 and are preferably structurally bonded between first skin 202and second skin 210. Also, if adhesive has been placed between adjacentportions of solid inserts 208 and large cell core sections 204, thensolid inserts 208 are joined with large cell core sections 204 and arepreferably structurally bonded with large cell core sections 204. In theillustrated embodiment, step 114 involves co-curing of solid inserts 208and large cell core sections 204 with first skin 202 and second skin210.

Once panel assembly 212 is cured, first skin 202, large cell coresections 204, solid inserts 208 and second skin 210 may be referred toas a large cell core stiffened panel 214 having solid inserts 208, whichmay also be represent by FIG. 4D. Importantly, the solid inserts 208provide a load path between first skin 202 and second skin 210. Inaddition, solid inserts 208 provide a large bonding surface for joiningwith first skin 202 and second skin 210 which reduces to potential forfirst skin 202 and/or second skin 210 from peeling away from large cellcore sections 204. In certain implementations, solid inserts 208 mayprovide a fluid barrier to limit fluid intrusion, such as fuel, waterand hydraulic fluid into large cell core sections 204.

Based upon the intended implementation for large cell core stiffenedpanel 214, additional finishing steps may be desired. In the illustratedembodiment wherein large cell core stiffened panel 214 is a lower wingskin assembly, step 116 includes machining or otherwise removing cutouts216 to provide fuel cell access, as best seen in FIG. 4E. In addition,as best seen in FIG. 4F, a plurality of holes 218 may be drilled throughfirst skin 202, solid inserts 208 and second skin 210 to provide joininglocations for coupling other airframe components to large cell corestiffened panel 214. For example, as best seen in FIG. 6A, a wing rib220 is joined to large cell core stiffened panel 214 such that wing rib220 is opposite a solid insert 208. As illustrated, wing rib 220 isjoined to large cell core stiffened panel 214 using a plurality ofaerospace fasteners 222 such as pins, screws, rivets or other suitablefastening means that extend through holes 218. Thus, the load applied tolarge cell core stiffened panel 214 by wing rib 220 during operation ofaircraft 10 is received at a solid insert 208, which is an ultra stiffload point within large cell core stiffened panel 214. Alternatively, asbest seen in FIG. 6B, an adhesive joint 226 is disposed between wing rib224 and large cell core stiffened panel 214 such that a lower surface ofwing rib 224 is structurally bonded to an upper surface of large cellcore stiffened panel 214. As illustrated, wing rib 224 is joined tolarge cell core stiffened panel 214 such that wing rib 224 is opposite asolid insert 208. Thus, the load applied to large cell core stiffenedpanel 214 by wing rib 224 during operation of aircraft 10 is received ata solid insert 208, which is an ultra stiff load point within large cellcore stiffened panel 214.

It is noted that the use of large cell core stiffened panels 214 asupper wing skin 58 and lower wing skin 56 of torque box structure 50advantageously provides a narrow profile for the large cell corestiffened wing skin that does not include any or require anyconventional stringers, thereby providing improved fuel bay clearance.In addition, the use of large cell core stiffened panels 214 as upperwing skin 58 and lower wing skin 56 improves the strength of torque boxstructure 50 including improved stiffness and torsional support duringthe shearing motion produced by proprotors 24 a, 24 b. The use of largecell core stiffened panels 214 results in a lower cost for manufacturingtorque box structure 50 due to reduced tooling and labor requirements, areduction in the part count and less intricate installation proceduressuch as a reduction in foam fillet installation. Further, the use oflarge cell core stiffened panels 214 as upper wing skin 58 and lowerwing skin 56 reduces the number of quality defects as compared to priorart implementations.

Referring next to FIGS. 7A-7B, therein is depicted a large cell corestiffened panel having solid inserts that is generally designated 300.Panel 300 includes a first skin 302 and a second skin 304 that arepreferably laminates formed from a plurality of material plies asdiscussed herein. A plurality of large cell core sections 306 are joinedbetween first skin 302 and second skin 304 and are preferablystructurally bonded between first skin 302 and second skin 304 asdiscussed herein. Large cell core sections 306 are preferably compositehoneycomb structures as discussed herein. A plurality of solid inserts308 are joined between first skin 302 and second skin 304 and arepreferably structurally bonded between first skin 302 and second skin304 as discussed herein. Solid inserts 308 are preferably compositeand/or compression molded inserts as discussed herein that havegenerally matching coefficients of thermal expansion with first skin 302and second skin 304. In the illustrated embodiment, solid inserts 308extend outwardly beyond the side surfaces of first skin 302 and secondskin 304. In addition, the outwardly extending ends of solid inserts 308are enlarged such that the upper and lower surfaces of the outwardlyextending ends of solid inserts 308 are generally flush with the lowersurface of first skin 302 and the upper surface of second skin 304. Eachof the outwardly extending ends of solid inserts 308 includes aplurality of holes 310 to provide joining locations for coupling othercomponents to large cell core stiffened panel 300.

Referring next to FIGS. 8A-8B, therein is depicted a large cell corestiffened panel having solid inserts that is generally designated 320.Panel 320 includes a first skin 322 and a second skin 324 that arepreferably laminates formed from a plurality of material plies asdiscussed herein. A large cell core section 326 is joined between firstskin 322 and second skin 324 and is preferably structurally bondedbetween first skin 322 and second skin 324 as discussed herein. Largecell core section 326 is preferably a composite honeycomb structure asdiscussed herein. A plurality of solid inserts 328 are joined betweenfirst skin 322 and second skin 324 and are preferably structurallybonded between first skin 322 and second skin 324 as discussed herein.Solid inserts 328 are preferably composite and/or compression moldedinserts as discussed herein that have generally matching coefficients ofthermal expansion with first skin 322 and second skin 324. In theillustrated embodiment, each of solid inserts 328 is positioned within acutout of large cell core section 326 and is fully surrounded by largecell core section 326. In addition, the side surfaces of each solidinsert 328 may be joined with and/or structurally bonded to the adjacentportions of large cell core section 326. First skin 322, solid inserts328 and second skin 324 include a plurality of holes 330 to providejoining locations for coupling other components to large cell corestiffened panel 320.

Even though the solid inserts of the present disclosure have beendepicted as symmetric solid inserts, it should be understood by thosehaving ordinary skill in the art that the solid inserts of the presentdisclosure could have any shape including non symmetric shapes. Forexample, as best seen in FIGS. 9A-9B, a large cell core stiffened panel340 includes non symmetric solid inserts. Panel 340 includes a firstskin 342 and a second skin 344 that are preferably laminates formed froma plurality of material plies as discussed herein. A large cell coresection 346 is joined between first skin 342 and second skin 344 and ispreferably structurally bonded between first skin 342 and second skin344 as discussed herein. Large cell core section 346 is preferably acomposite honeycomb structure as discussed herein. Solid inserts 348,350 are joined between first skin 342 and second skin 344 and arepreferably structurally bonded between first skin 342 and second skin344 as discussed herein. Solid inserts 348, 350 are preferably compositeand/or compression molded inserts as discussed herein that havegenerally matching coefficients of thermal expansion with first skin 342and second skin 344. In the illustrated embodiment, each of solidinserts 348, 350 is positioned within a cutout of large cell coresection 346 and is fully surrounded by large cell core section 346 suchthat the outer surfaces of solid inserts 348, 350 may be joined withand/or structurally bonded to the adjacent portions of large cell coresection 346. Solid insert 348 includes a central void 352 and solidinsert 350 includes a central void 354 that enable reduction in theweight of solid inserts 348, 350. First skin 342, solid insert 348 andsecond skin 344 include a plurality of holes 356 to provide joininglocations for coupling other components to large cell core stiffenedpanel 340. Likewise, first skin 342, solid insert 350 and second skin344 include a plurality of holes 358 to provide joining locations forcoupling other components to large cell core stiffened panel 340.

Referring next to FIGS. 10A-10B, therein is depicted a large cell corestiffened panel having solid inserts that is generally designated 360.Panel 360 includes a first skin 362 and a second skin 364 that arepreferably laminates formed from a plurality of material plies asdiscussed herein. A large cell core section 366 is joined between firstskin 362 and second skin 364 and is preferably structurally bondedbetween first skin 362 and second skin 364 as discussed herein. Largecell core section 366 is preferably a composite honeycomb structure asdiscussed herein. A plurality of solid inserts 368 are joined betweenfirst skin 362 and second skin 364 and are preferably structurallybonded between first skin 362 and second skin 364 as discussed herein.Solid inserts 368 are preferably composite and/or compression moldedinserts as discussed herein that have generally matching coefficients ofthermal expansion with first skin 362 and second skin 364. In theillustrated embodiment, each of solid inserts 368 is positioned within acutout of large cell core section 366 and is fully surrounded by largecell core section 366. In addition, the side surfaces of each solidinsert 368 may be joined with and/or structurally bonded to the adjacentportions of large cell core section 366. In the illustrated embodiment,solid insert 368 each include a pass through opening 370 which may bepreformed in solid inserts 368 or formed together with correspondingportions of pass through openings 370 in first skin 362 and second skin364. Pass through openings 370 may provide locations for fluid lines,electrical lines, data cables or the like to pass through panel 360.

Referring next to FIGS. 11A-11B, therein is depicted a large cell corestiffened panel having solid inserts that is generally designated 380.Panel 380 includes a first skin 382 and a second skin 384 that arepreferably laminates formed from a plurality of material plies asdiscussed herein. A large cell core section 386 is joined between firstskin 382 and second skin 384 and is preferably structurally bondedbetween first skin 382 and second skin 384 as discussed herein. Largecell core section 386 is preferably a composite honeycomb structure asdiscussed herein. A plurality of solid inserts 388 are joined to firstskin 382 and second skin 384 and are preferably structurally bonded tofirst skin 382 and second skin 384 as discussed herein. Solid inserts388 are preferably composite and/or compression molded inserts asdiscussed herein that have generally matching coefficients of thermalexpansion with first skin 382 and second skin 384. In the illustratedembodiment, each of solid inserts 388 is positioned within a cutout oflarge cell core section 386 and is fully surrounded by large cell coresection 386. In addition, the side surfaces of each solid insert 388 maybe joined with and/or structurally bonded to the adjacent portions oflarge cell core section 386. In the illustrated embodiment, solid insert388 each include a cylindrical extension that passes through a precutopening 390 in second skin 384 such that the side surfaces of eachcylindrical extension may be joined with and/or structurally bonded tothe adjacent side surfaces of openings 390. The cylindrical extensionsmay provide joining locations for coupling other components to largecell core stiffened panel 380.

The foregoing description of embodiments of the disclosure has beenpresented for purposes of illustration and description. It is notintended to be exhaustive or to limit the disclosure to the precise formdisclosed, and modifications and variations are possible in light of theabove teachings or may be acquired from practice of the disclosure. Theembodiments were chosen and described in order to explain the principalsof the disclosure and its practical application to enable one skilled inthe art to utilize the disclosure in various embodiments and withvarious modifications as are suited to the particular use contemplated.Other substitutions, modifications, changes and omissions may be made inthe design, operating conditions and arrangement of the embodimentswithout departing from the scope of the present disclosure. Suchmodifications and combinations of the illustrative embodiments as wellas other embodiments will be apparent to persons skilled in the art uponreference to the description. It is, therefore, intended that theappended claims encompass any such modifications or embodiments.

What is claimed is:
 1. An airframe assembly for an aircraft comprising:a first airframe member having a first skin, a second skin, a large cellcore joined between the first and second skins and a solid insert havinga side surface, the solid insert joined between the first and secondskins such that at least a portion of the side surface is adjacent tothe large cell core, the first airframe member having a first set ofopenings extending through the first skin, the solid insert and thesecond skin; a second airframe member having a second set of openingsoperable to be aligned with the first set of openings of the firstairframe member; and a plurality of fasteners each extending through oneof the openings of the first set of openings and one of the openings ofthe second set of openings securably coupling the first airframe memberto the second airframe member.
 2. The airframe assembly as recited inclaim 1 wherein the large cell core is joined to the adjacent portion ofthe side surface of the solid insert.
 3. The airframe assembly asrecited in claim 1 wherein the large cell core is selected from thegroup consisting of large cell composite cores and large cell carboncores.
 4. The airframe assembly as recited in claim 1 wherein the firstskin, the second skin and the solid insert further comprise a materialselected from the group consisting of composite materials and carboncomposite materials.
 5. The airframe assembly as recited in claim 1wherein the first skin, the second skin and the solid insert havegenerally matching coefficients of thermal expansion.
 6. The airframeassembly as recited in claim 1 wherein the large cell core and the solidinsert are structurally bonded between the first and second skins. 7.The airframe assembly as recited in claim 1 wherein the large cell coreis structurally bonded to the adjacent portion of the side surface ofthe solid insert.
 8. A wing assembly for an aircraft comprising: a wingskin having an inner skin member, an outer skin member, a large cellcore joined between the inner and outer skin members and a solid inserthaving a side surface, the solid insert joined between the inner andouter skin members such that at least a portion of the side surface isadjacent to the large cell core, the wing skin having a first set ofopenings extending through the inner skin member, the solid inset andthe outer skin member; a rib having a second set of openings operable tobe aligned with the first set of openings of the wing skin; and aplurality of fasteners each extending through one of the openings of thefirst set of openings and one of the openings of the second set ofopenings securably coupling the rib to the wing skin.
 9. The wingassembly as recited in claim 8 wherein the large cell core is joined tothe adjacent portion of the side surface of the solid insert.
 10. Thewing assembly as recited in claim 8 wherein the large cell core isselected from the group consisting of large cell composite cores andlarge cell carbon cores.
 11. The wing assembly as recited in claim 8wherein the inner skin member, the outer skin member and the solidinsert further comprise a material selected from the group consisting ofcomposite materials and carbon composite materials.
 12. The wingassembly as recited in claim 8 wherein the inner skin member, the outerskin member and the solid insert have generally matching coefficients ofthermal expansion.
 13. The wing assembly as recited in claim 8 whereinthe large cell core and the solid insert are structurally bonded betweenthe inner and outer skin members.
 14. The wing assembly as recited inclaim 8 wherein the large cell core is structurally bonded to theadjacent portion of the side surface of the solid insert.
 15. Anaircraft comprising: a first airframe member having a first skin, asecond skin, a large cell core joined between the first and second skinsand a solid insert having a side surface, the solid insert joinedbetween the first and second skins such that at least a portion of theside surface is adjacent to the large cell core, the first airframemember having a first set of openings extending through the first skin,the solid insert and the second skin; a second airframe member having asecond set of openings operable to be aligned with the first set ofopenings of the first airframe member; and a plurality of fasteners eachextending through one of the openings of the first set of openings andone of the openings of the second set of openings securably coupling thefirst airframe member to the second airframe member.
 16. The aircraft asrecited in claim 15 wherein the large cell core is joined to theadjacent portion of the side surface of the solid insert.
 17. Theaircraft as recited in claim 15 wherein the large cell core is selectedfrom the group consisting of large cell composite cores and large cellcarbon cores.
 18. The aircraft as recited in claim 15 wherein the firstskin, the second skin and the solid insert further comprise a materialselected from the group consisting of composite materials and carboncomposite materials.
 19. The aircraft as recited in claim 15 wherein thefirst skin, the second skin and the solid insert have generally matchingcoefficients of thermal expansion.
 20. The aircraft as recited in claim15 wherein the large cell core and the solid insert are structurallybonded between the first and second skins.
 21. The aircraft as recitedin claim 15 wherein the large cell core is structurally bonded to theadjacent portion of the side surface of the solid insert.